The invention relates to an inner ring for a guide vane assembly for mounting adjustable guide vanes of a turbomachine, of a compressor stage or turbine stage of a gas turbine, a guide vane assembly for a turbomachine having an inner ring, a turbomachine having an inner ring, and a method for producing an inner ring, wherein the inner ring has a plurality of guide vane bearing mounts, which are each arranged spaced apart in the peripheral direction, recesses, which are each formed for receiving a bearing element, a bearing journal, of a guide vane, wherein, between at least two adjacent guide vane bearing mounts, the inner ring has at least one depression with a wall thickness in a radial direction that is reduced in comparison to a region that abuts the depression and is outside of the depression.

A component is provided for a gas turbine engine. This component includes a monolithic body configured with an internal passage. The internal passage extends along a centerline within the monolithic body. The internal passage has a cross-sectional geometry perpendicular to the centerline. The cross-sectional geometry of at least a first portion of the internal passage has a teardrop shape.

A component is provided for a gas turbine engine. This component includes a monolithic body configured with an internal passage. The internal passage extends along a centerline within the monolithic body. The internal passage has a cross-sectional geometry perpendicular to the centerline. The cross-sectional geometry of at least a first portion of the internal passage has a teardrop shape.

A turbo machine includes a casing in which fluid flows; turbine blades arranged side by side in a circumferential direction with respect to a rotational shaft rotatably provided in the casing; a ring segment forming an inner surface of the casing; a seal fin provided as an extension at a tip end part of each turbine blade and facing the ring segment; and a seal member through which the turbo machine faces the seal fin, the seal member having a first inner surface that accepts contact of the seal fin. The ring segment has a second inner surface incrementally enlarging toward a downstream side in an axial direction of the rotational shaft. The first and second inner surfaces are connected with each other without a step between the inner surfaces on the downstream side of the first inner surface in the axial direction.

An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

A waste-gate valve is for opening and closing a bypass passage which bypasses an exhaust turbine of a turbocharger, and includes: a valve body disposed in the bypass passage; an open-close lever having a first insertion hole into which a valve shaft of the valve body is inserted, and being configured to open and close the bypass passage by moving the valve body; and a washer having a second insertion hole which is positioned closer to a tip of the valve shaft than the first insertion hole and into which the valve shaft is inserted, the washer being fixed to the valve shaft. The washer has a bend portion bended along an outer shape of the open-close lever.

An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

A fastener cover assembly is disclosed. The cover assembly includes a turbomachinery component having an outer surface exposed to an air flow path; at least one recess formed in the outer surface; a fastener contained in the at least one recess; and a fastener cover secured in the at least one recess, the cover having an outer surface substantially flush with the outer surface of the component to minimize air flow disruption.

An exhaust system or nozzle for use in a gas turbine engine is disclosed herein. The exhaust system is adapted to adjust various streams of pressurized air produced by the gas turbine engine to control operation of the gas turbine engine.

A gas turbine engine component includes a structure having a surface configured to be exposed to a hot working fluid. The surface includes a recessed pocket that is circumscribed by an overhang. At least one cooling groove is provided by the overhang.